1. Field of the Invention
This invention relates to gas turbine engine turbine vane repair and, more particularly, to turbine vane with gap brazed repairs having localized compressive residual stresses imparted by laser shock peening.
2. Description of Related Art
A gas turbine engine includes a compressor section, a combustion section and a turbine section. Disposed within the turbine section are alternating annular stages of circumferentially disposed moving blades and stationary vanes. The rows or stages of vanes and blades are concentrically located about a center-line axis of the gas turbine engine. The blades are mounted on a disk which rotates about its central axis. Hot combustion gases exit the combustor and pass through the turbine section, whereupon the blades which are mounted on the disks rotatably drive a shaft, thus providing shaft work for driving the fan and compressor sections and auxiliary systems. Higher gas temperatures allow more work to be extracted from the gases in the turbine section, thus increasing the overall efficiency of the gas turbine engine.
The stationary vanes disposed between the stages of moving blades stabilize and direct the gas flow from one stage of rotating turbine blades to the next stage of rotating turbine blades. The stabilization and turning of the gas flow optimizes the amount of work extracted from the hot gases in the turbine section. It is very important to the efficient operation of the turbine vanes and engine to maintain the structural integrity of the nozzle and, in particular, the nozzle flow areas which are the spaces between adjacent vanes in a vane stage. Cobalt and nickel-base superalloy materials have been developed to provide mechanical strength at high temperatures so that the operating temperature capability of the turbine section is increased over the operating temperatures of prior designs. It is known to cast engine vanes from superalloys, for example, a nickel-base alloy or a cobalt-base alloy. Nickel-base superalloys are frequently primarily strengthened by precipitation of a gamma prime phase, Ni3 (Al, Ti) when used in gas turbine engines and, in particular, in turbine vanes. In addition, the casting of turbine vanes and blades is frequently performed so as to produce a directionally solidified part, with grains aligned parallel to the axis of the blade or vane or a single crystal part, with no grain boundaries.
In service, deterioration of the vane surface occurs due to oxidation, thermal fatigue cracking and metal erosion caused by abrasives and corrosives in the flowing gas stream. In addition, the high gas pressures at high temperature cause distortion of the vanes, thereby, enlarging the nozzle area with a consequent loss in turbine efficiency. During periodic engine overhauls, the vanes are inspected for physical damage and measured to determine the degree of flow area change and the effect on nozzle flow area. Before these vanes can be returned to the engine physical damage is usually repaired or the vanes replaced.
Several methods exist for repairing the worn or damaged vanes and for returning the nozzle gas flow area to acceptable dimensions within predetermined tolerances. Repair methods include, for example, conventional fusion welding, plasma spray as described in U.S. Pat. No. 4,878,953, and the use of a tape or slurry material containing a mixture of a binder and a metal alloy powder which is compatible with the substrate alloy. U.S. Pat. No. 4,878,953 provides an excellent source of background information related to methods for refurbishing cast gas turbine engine components and particularly for components made with nickel-base and cobalt-base superalloys for use in the hot sections of gas turbine engines and, more particularly, for components exposed to high temperature operating conditions.
Cracking resulting from high thermal stresses and oxidation are often repaired with a wide gap braze alloy. One particular exemplary repair method provides for filling cracks which can be as long as approximately 1.5" and as wide as approximately 0.050" in order to refurbish vanes in this manner. Trailing edges in particular are subjected to the most severe conditions, so much so that vanes can be returned from the field with missing portions of trailing edges and in many cases vanes are considered unrepairable. The HPT vane materials for the latest engines require high strength oxidation resistant single crystal and directionally solidified cast alloys. The current wide gap braze repair system for repairing vanes made of these alloys and by single crystal processes do not restore all of these properties in the repaired areas of the vane. Further, life enhancement of the repair and vane at a low cost is therefore highly desirable. Vane replacement is very costly and there is always a need for stronger, longer lasting, and more cost effective vane repairs. It is very desirable to provide life extension of the vane and vane repair which would further approach single crystal properties, offer wider repairability limits, and make available repairs on hardware currently considered unrepairable because of excessive damage.
The present invention is directed towards this end and provides a wide gap braze repair, particularly, for a gas turbine engine vane, with regions of deep compressive residual stresses imparted by laser shock peening the area over the repair.
The region of deep compressive residual stresses imparted by laser shock peening of the present invention is not to be confused with a surface layer zone of a work piece that contains locally bounded compressive residual stresses that are induced by a hardening operation using a laser beam to locally heat and thereby harden the work piece such as that which is disclosed in U.S. Pat. No. 5,235,838, entitled "Method and Apparatus for Truing or Straightening Out of True Work Pieces". The present invention uses multiple radiation pulses from high power pulsed lasers to produce shock waves on the surface of a work piece similar to methods disclosed in U.S. Pat. No. 3,850,698, entitled "Altering Material Properties"; U.S. Pat. No. 4,401,477, entitled "Laser Shock Processing"; and U.S. Pat. No. 5,131,957, entitled "Material Properties". Laser peening, as understood in the art and as used herein, means utilizing a laser beam from a laser beam source to produce a strong localized compressive force on a portion of a surface. Laser peening has been utilized to create a compressively stressed protection layer at the outer surface of a workpiece, which is known to considerably increase the resistance of the workpiece to fatigue failure as disclosed in U.S. Pat. No. 4,937,421, entitled "Laser Peening System and Method". However, the prior art does not disclose wide gap braze repairs of the type claimed by the present patent nor the methods of how to produce them. It is to this end that the present invention is directed.